The present invention relates to a new Satellite Configuration for geostationary missions.
Conventional satellite configurations consist of self contained units which include all service function elements that are required by the specific mission payload as well as the payload itself. In view of satellite reliability and service availability considerations sufficient on-board redundancy needs to be achieved to meet the life time mission requirements. Due to the mass limitation of some launch vehicles, the redundancy requirement leads to the adoption of satellite systems that require, for example for a 10 year service life, the procurement a number of individual flight satellites and their launch with a primary and a spare operational satellite always in orbit, and at least one spare satellite on the ground.
Typical disadvantages of the conventional multisatellite system are:
(a) High satellite and launch costs due to the number satellites required,
(b) The satellite systems are complex due to the effort of maximizing single satellite mission usefulness constrained by mass limitations of the launch vehicle,
(c) There is a high orbit occupancy, for example of geostationary orbit positions, for a single mission service,
(d) High operational costs due to the monitoring and control of two in-orbit satellites for each system,
(e) Danger of interference between the growing number of satellites around desired or allocated orbit positions.
The problem to be solved is to provide a geostationary satellite configuration which at the same time meets mission requirements and overcomes the mass limitations of the launch vehicle.